All fighter aircraft projects are contractually required to perform a full-scale spin demonstration program. The purpose of the flight test program is to identify the types of spins that could be encountered inadvertently during future operational use of the aircraft, and the control techniques that are required to return the aircraft to the normal flight regime. Other purposes for these programs could include the demonstration of aerodynamic configurations that are spin resistant, automatic spin avoidance and/or recovery techniques, etc. In any event, spin demonstration aircraft are required now and will be required in the future, and these aircraft must be equipped with an emergency recovery system that is guaranteed to terminate any otherwise unrecoverable spin mode that might be encountered.
One spin mode to which modern fighter-type aircraft are susceptible is the flat spin, wherein the aircraft exhibit "spinning top" motions. This type of spin usually has a high rate of rotation, an angle of attack between 70.degree. and 90.degree., and effectively no spin radius, the aircraft spinning about an axis that passes through or near the center-of-gravity of the aircraft. Because an aircraft usually cannot recover from a developed flat spin through manipulation of the available aerodynamic controls, it is the spin which pilots fear most.
To maintain a flat spin or any other type of spin, the aircraft must balance the nose-down aerodynamic pitching moment with an opposing gyroscopic pitching moment. The magnitude of the aerodynamic pitching moment is a function of the aircraft configuration, dynamic pressure and angle of attack (usually increasing progressively up to 90.degree. angle of attack). The magnitude of the gyroscopic pitching moment is a function of the mass distribution and the product of the roll and yaw rates generated about the aircraft body axis. In a flat spin, the yaw rate is considerably greater than the roll rate. The yaw rate required for spin equilibrium is determined by the magnitude of the aerodynamic pitching moment and the aircraft mass distribution. The other requirement for spin equilibrium is that the aerodynamic yawing moment about the body yaw axis be zero (actually very slightly propelling, i.e., pro-spin) at this yaw rate. Obtaining a flat spin requires, therefore, that a propelling aerodynamic yawing moment be generated at yaw rates below that required for balancing the aerodynamic pitching moment and that the magnitude of this yawing moment decrease (approaching a zero value) as the required yaw rate is attained. If a damping (anti-spin) yawing moment is generated below and at the required yaw rate the flat spin cannot be maintained.
Emergency recovery systems used to date to generate an anti-spin yawing moment are complex, and usually incorporate a tail chute which is extremely inefficient when installed on modern aircraft that spin flat. In some instances, the chute size which is required for a particular type aircraft becomes impracticably large. In addition, the length of the riser line that attaches the tail chute to the aircraft is critical. If the riser line length is too short, the chute tends to collapse in the low dynamic pressure and reversed flow field that exists above the aircraft. If the riser line length is too long, the chute trails the aircraft at an angle which results in a nose-down pitching moment but no anti-spin yawing moment. Even the optmum riser line length results in a chute trail angle that contributes only a small antispin yawing moment. To compensate for the small anti-spin yawing moment, large parachutes are used. However, the use of large chutes results in off-design loads on the aircraft, which necessitates extensive internal and external reinforcement of the fuselage. In some cases, the reinforcement of the fuselage incurs changes in the mass distribution and external shape of the spin demonstration aircraft which jeopardizes the applicability of the results obtained from the testing.
National Aeronautics and Space Administration has published a survey of spin-recovery devices and techniques, including rockets and wing-tipped mounted parachutes, but with particular emphasis on approaches in the design of tail-mounted spin-recovery parachute systems, including a compilation of design considerations applicable to spin-recovery parachute systems. The survey is published as NASA Technical Note D-6866, entitled "Summary of Design Considerations for Airplane Spin-Recovery Parachute Systems", by Sanger M. Burk, Jr., published in August, 1972. The disclosure of that publication is herein incorporated by reference, since many forms of apparatus disclosed therein are usable in my invention, even though my invention involves a basically contrary approach. The publication is also noteworthy in that it evidences the complexity of the problems involved and the great deal of effort which has gone toward their solution.
Reference is here made also to my copending application, Ser. No. 570,505, filed Apr. 22, 1975, for Aerodynamic Spin Control Device for Aircrafts, wherein I disclose and claim a different solution to the same general problem, involving the use of doors or like devices mounted in the forebody of the aircraft so as to be deployable outwardly under spin conditions so as to change the fuselage yawing moment from pro-spin to anti-spin by changing the flow field over the fuselage forebody at high angles of attack, by effectively changing the geometric characteristics of the forebody.
As indicated in the NASA publication referred to above, emergency spin recovery parachute systems have been traditionally attached to the aft end of the fuselage, although wing-tip-mounted parachutes have been given consideration, with notable lack of success. The tail-mounted technique resulted in a reasonably efficient recovery system for aircraft developed during the 1930-l950 time period. Since then, however, radically different airframe-propulsion configurations have evolved for which I consider this system to be unsuited. The efficiency of this technique progressively deteriorated over the years as the total plan form area increased, the distance between the center-of-gravity location and aft end decreased, and a greater proportion of the mass became concentrated in the fuselage. As the efficiency eroded, the complexity, size, and structural attachment load of the recovery installation escalated correspondingly. For some recent designs, the required structural modification costs have reached an absurd level.
The major components of a conventional emergency spin recovery system installed at the aft end of the fuselage are: a deployment motor or gun, pilot chute, bridle line, deployment bag, riser line, ring slot or ribbon-type main chute, chute inflating device, and a chute tiedown and release mechanism. This emergency system is usually deployed during a flat spin, which is the most difficult spin to terminate, this being the case since the spin rate is the highest achievable while the aerodynamic controls required for recovery are operating at their lowest effectiveness. Also, the use of the controls may be lost completely on some aircraft after an engine flame-out is experienced. Unfortunately, when the need is greatest for the emergency system it is grossly inefficient since by virtue of its location on modern aircraft, which spin about a vertical axis passing through the center of gravity, only a small portion of the chute load produces the anti-spin yawing moment required for overcoming the spin rotation, whereas most of the chute load applies a nose-down pitching moment which generates a gyroscopic pro-spin yawing moment on current aircraft which have more of their mass concentrated in the fuselage than along the wing. Obviously, attachment of the recovery chute to the aft end of a high performance aircraft does not employ the laws of nature in an optimum manner but attempts to oppose them instead.
The inefficiency of the tail-mounted chute location dictates the need for a large chute, which incurs a need for a major structural beef-up (accomplished internally and/or externally), an undesirable change in external lines, an undesirable increase in the inertia and change in the mass distribution, a bulky installation requiring special packing, inspection, etc. to keep size to a minimum, and a need for separating the chute from the aircraft before regaining normal flight. It should be noted that in one instance a 48 feet diameter chute attached to a 130 feet riser line was found to be inadequate. Also, the chute location is inefficient on other grounds which dictate additional severe design requirements. For example, the chute and riser line must be protected from heat before, during and after deployment of the system. The system must be designed to ensure minimum contact with aircraft structure, that is, avoid fouling of pilot chute and cuts and abrasion on riser line. The chute must trail in a stable position (not oscillate) to ensure an anti-spin yawing moment. This dictates the need for a high porosity type chute which requires an inflation device. The riser line length is critical, therefore, hopefully chosen properly since the main chute will collapse in the low q (dynamic pressure) reversed flow field above the aircraft if the line is too short. If it is too long, no anti-spin yawing moment is generated.
The consequences of the described inefficiencies are high cost, long lay-ups of the test vehicle, system complexity, and a need for extensive (yet unsatisfactory) checkout system tests for evaluating a configuration which no longer represents the production vehicle. In addition, chute redundancy is not feasible, nor can installation on a production aircraft be considered.
My invention incorporates a very small pilot type chute or chutes which can be stored on or in, and attached to, the nose of the aircraft, and which can be deployed therefrom. This device is highly effective because the chute location in operation takes advantage of the aircraft characteristics inherent in a high performance aircraft. For example, at deployment the total chute load applies a side force through a large moment arm and, consequently, a significant aerodynamic anti-spin yawing moment. As the rotation rate decreases and the chute realigns itself with the relative wind, a portion of the chute load applies a normal force and therefore a nose-up pitching moment which creates a gyroscopic anti-spin yawing moment. The chute operates in a relatively free-stream environment.
Since the nose chute is extremely efficient, the chute size and corresponding load are very small, Therefore, no structural beef-up or external modifications are required, and no weight, inertia or mass distribution change need be incurred. This location also allows the use of an unstable chute. A low porosity chute can therefore be employed which in turn minimizes the required size and volume and needs no inflation device. The efficiency can be further enhanced through the use of sequenced chutes which are preprogrammed in an optimal fashion. A sequenced chute system, therefore, employs chutes which are smaller in size than would be required for a single chute system.
In keeping with my invention, the parachute is deployed upon entry into a condition of incipient or developed spin, which deployment may be by the pilot upon entry into the spin, or in the incipient spin phase in response to a predetermined angle of attack and yaw rate sensed by a conventional air data computer, or automatically under the control of a spin condition sensing means carried by the aircraft. Although the chute size is variable, as one typical measure of size it may be said that the size of the chute, when deployed, is such as to have a negligible influence on the free falling rate of sink of the aircraft. For most aircraft, a maximum chute diameter of less than 10 feet is appropriate. When mounted or coupled forward of the cockpit, the dimensions of the chute should be such that the deployed chute can be collapsed against the side of the fuselage during a spin recovery dive without interfering with pilot visibility. The chute itself, or the chutes, may be considered as conventional, and may be mounted externally on the fuselage in a conventional chute bag, or may be housed in a special housing inside the fuselage and deployed therefrom in what may be considered a conventional manner.
I am aware that there have been previously proposed various arrangements of parachutes in the nose portion of an aircraft, although for completely different purposes, these purposes being typically to hopefully lower an aircraft gently to the ground after a power failure. A typical example is found in Frost U.S. Pat. No. 2,673,051, issued Mar. 23, 1954, wherein a large supporting parachute is housed in the nose portion of a propeller driven VTOL aircraft so as to be deployable during vertical takeoff or landing in the event of a power failure. As a supporting parachute, the parachute would have to be necessarily quite large, and there is no relationship between the concept of this arrangement and my invention. A further typical example is found in Krahel U.S. Pat. No. 2,352,721, issued July 4, 1944, wherein large parachutes are coupled to the nose and other portions of the aircraft so as theoretically to lower the aircraft to the ground and prevent a crash. British Pat. No. 1,057,362, published Feb. 1, 1967 (McMahon and Brown) discloses a parachute arrangement for lowering a rocket to the ground, the parachute being coupled to the forebody of the rocket and to the tail portion. These examples are conceptually remote from my invention, but they constitute the only instances known to me wherein a parachute is coupled to the forebody of an aircraft, regardless of the purpose.
Other features and advantages of my invention will be set forth in or apparent from the following detailed description of presently preferred embodiments, taken in conjunction with the attended drawings.